1. Field of the Invention
The present invention relates to a state control device of a moving body and a state control method of a moving body, particularly to a state control device of a moving body and a state control method of a moving body in a spacecraft.
2. Description of Related Art
Conventionally, there have been proposed many methods in a state control system including, for example, attitude control or the like in a moving body including a spacecraft.
However, according to a spacecraft such as an artificial satellite or the like, a request of mount mission in recent years has become severe in accuracy and accordingly, for example, the following promotion in function is needed.
1) To promote accuracy in attitude control in respect of a target attitude value of a satellite. (reduce an absolute value of error) PA1 2) To promote attitude stability in respect of the target attitude value of a satellite. (restrain rate variation at low frequencies, in other words, realize a satellite which is stationary even with disturbance) PA1 3) To promote response in respect of the target attitude value of a satellite. PA1 4) To provide high directional capacity in a satellite configuration having a large-sized and flexible solar cell paddle, a drive antenna having large angular momentum or other pointing devices. PA1 5) To achieve shortening of an adjustment time period and promotion of design efficiency by making interface with other subsystem constituting a disturbance source exemplified as described above as simple as possible. PA1 6) With regard to the above-described 5), to be able to further promote removal of an attitude error which cannot be removed by a feedforward system of an attitude control system based on a drive signal (angle/angular velocity or the like) of an antenna which has been carried out generally.
In the meantime, as a general attitude control system of a spacecraft which has been conventionally known, as shown by FIG. 7, information of rotational angle information, rotational angular velocity information or the like of a spacecraft 1 which is outputted from a navigation dynamics 2 of the satellite inherently provided to the spacecraft 1 per se, is detected by an attitude sensor 3, the information is inputted to an attitude control system 4 for executing, for example, a PID (Proportional ,Integral, Differential) control and the PID control system 4 generates a control command for driving wheels and the like to thereby rotate the wheels by predetermined amounts by which the attitude of the spacecraft 1 is controlled.
In such a case, various disturbance is applied to the spacecraft 1 and therefore, it is general to constitute the control system by previously anticipating the amount of disturbance.
For example, Japanese Unexamined Patent Publication No. JP-A-8-188199 discloses a technology of installing estimated disturbance removing signal generating means 6 in which the magnitude of disturbance applied on the main body of a satellite is estimated, a disturbance removing signal for removing the disturbance noise is generated and an estimated disturbance removing signal is generated along with an attitude control signal provided from the attitude control system.
Meanwhile, according to such an attitude control system of a spacecraft, there causes enormous restriction in constructing a high accuracy attitude control system by being significantly influenced by accuracy in parameters of a drive system, that is, factors of mass characteristic, angle detection accuracy, calculation accuracy of angular velocity, calculation accuracy of angular acceleration, communication delay, jitter component of communication time period accompanied by synchronous processings among different computers and so on.
For example, Japanese Unexamined Patent Publication No. JP-A-3-125699 discloses a method in which in the attitude control of a spacecraft, a feedforward compensation signal is used and high frequency components are removed from the feedforward compensation signal.
According to such a publicly-known example, in carrying out the feedforward calculation, when angular velocity derived from a quasi differential value (difference) is calculated from angle information of a drive system at a control period and angular momentum or angular acceleration of a further higher order, that is, disturbance torque is calculated, time-sequential data of angular velocity obtained from the drive system by a low pass filter, is smoothed to remove noise components and low frequency components are separated.
Therefore, according to such a publicly-known example, although the noise components caused in calculating the quasi differential value are eliminated to some degree, instead, a delay is caused by time constant of the low pass filter, a function of "adding compensation torque having a reverse phase simultaneously with the drive system" which is most important in the feedforward control, is sacrificed and a restriction seems to be caused in realizing promotion of accuracy in attitude control or attitude stability in which the above-described requested value has become severe.
Further, when the accuracy is intended to improve, a computer for an attitude system needs to calculate in synchronism with a computer for mounting the drive system, communication interruption processings are needed or tolerance of mass characteristic of the drive system needs to prescribe to an extreme which may significantly enhance function, adjustment time period and design cost.
That is, although disturbance per se cannot be actually measured and therefore, the disturbance needs to estimate by some method, when the estimation is intended to be accurate, the system per se becomes complicated which amounts to an increase in cost and further, a plurality of computers mounted on the spacecraft 1 are operated asynchronously to each other and accordingly, timing of the processing of the disturbance information needs to match which makes the system further complicated.
Further, in the spacecraft 1 of recent times, when the structure of an antenna 8 or a solar cell panel 7 is large-sized and a flexible structure is adopted, antenna maneuver disturbance caused by the antenna occurs, movement and vibration of the antenna or the solar cell panel is enhanced and constitutes low frequencies and therefore, reaction which is disturbance imposed on the spacecraft 1 becomes significant.
Therefore, with regard to attitude control against such reaction, the conventional control system amounts to an increase in cost and complication of the system and accurate control becomes difficult.
Otherwise, for example, according to Japanese Unexamined Patent Publication No. JP-A-62-125998, there is described an attitude control method of a spacecraft in which a feedback signal storing unit is installed in a conventional closed loop control, a change in a feedback signal is stored to the storing unit and error compensation command is determined such that attitude error is reduced based on the data and according to Japanese Unexamined Patent Publication No. JP-A-8-282598, there is disclosed a technology in an attitude control device in an artificial satellite for installing torque compensating means for generating torque necessary for canceling angular momentum from an actuator and drive controlling means which is constituted such that drive of a movable unit is retarded more than generation of the torque, however, any of them are insufficient for resolving the above-described problem.